During operation of a gas turbine, the components, inter alia turbine blades, blade carriers and blade platforms, are exposed to hot gas temperatures of over 1400° C. The turbine components in the hot gas region are provided with a metallic oxidation layer and also often with a ceramic thermal-insulating layer, also called “thermal barrier coating” (TBC), such that the material can withstand these temperatures. In addition, the turbine components are cooled by means of cooling air, which flows through a complex system of cooling passages. The cooling passages open out to the surface of the components and often have in the region of the surface a specific three-dimensional shape which ensures film cooling of the component surface by the cooling air. In order to ensure reliable operation of the turbine, certain components, in particular turbine blades, are removed and replaced after an operating interval, for example of 20000 hours. In this case, it is commercially advantageous to recondition turbine blades and use them again for an operating interval. During reconditioning, the protective layers are typically removed and applied again, the outlet openings of the cooling passages being obstructed by the protective layer material applied again. In order to achieve full cooling capacity, the cooling passages must be opened again, in which case the geometry of the original cooling air holes should preferably be restored as accurately as possible.
In a turbine blade used nowadays, several hundred individual cooling passages, for example, are distributed at the surfaces, and these cooling passages can be divided, for example, into about twenty different basic types of passage. The basic types differ, for example, in their orientation relative to the surface of the turbine component, in the size of their cross section or in the angles of spread in the outlet region of the cooling passages.
Recoating is in many cases carried out by plasma spraying and with a spraying direction that is as perpendicular as possible to the component surface. The obstruction of the individual cooling passages therefore varies greatly, depending on the orientation and cross-sectional size of the passages and on their position relative to the contour of the component. In addition, the restoration of the original passage geometry is made more difficult by the large tolerances of the plasma spraying process by virtue of the fact that the thickness of the obstructing material can be predicted only to a limited extent.
EP 1 510 283 discloses a method of restoring cooling passages in turbine blades. First of all a local reference coordinate system is prepared by means of characteristic features of the component. In an automated scanning process, the three-dimensional positions and orientations of each cooling passage which are related to a local reference coordinate system are determined. The position and orientation data are then used for removing the obstructing coating material and for restoring the original cooling passage. The unwanted material is removed by vaporization by means of a pulsed laser (laser ablation).